The present invention relates generally to gas turbine engines, and, more specifically, to blade damping therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which powers the compressor. Additional energy is extracted in a low pressure turbine (LPT) for powering a fan in a turbofan aircraft engine application, or for powering an output drive shaft for marine and industrial (M&I) applications.
Each turbine stage includes a row of turbine rotor blades extending radially outwardly from a supporting rotor disk. During operation, the turbine blades extract energy from the hot combustion gases and are subject to considerable pressure and centrifugal loads, and operate at elevated temperature.
The blades are typically hollow and include internal cooling circuits through which a portion of pressurized air bled from the compressor is circulated for cooling the individual blades against the heat loads from the combustion gases.
The turbine blades typically increase in size and length in the successive stages from the combustor for maximizing efficiency of energy extraction as the pressure in the combustion gases decreases in the downstream direction. The turbine blades are subject to vibratory excitation forces due to the aerodynamic and centrifugal loads and behave differently due to the different sizes and configurations of the blades, with different modes of vibration occurring at different resonant frequencies.
Vibration damping may be effected where desired using under platform vibration dampers, or internal vibration dampers installed inside the individual blades.
Internal vibration dampers typically extend the length of the turbine blade and are mounted inside the supporting dovetail and are cantilevered freely inside the airfoil.
Each damper is typically a slender rod having a small lean so that centrifugal forces load the damper radially outwardly against corresponding internal supporting lands inside the airfoil. Frictional vibration between the damper and the airfoil dissipates excitation forces and effectively dampens blade vibration.
However, frictional damping is subject to wear between the damper and the airfoil, and should be minimized. Yet, the damper itself is subject to substantial centrifugal loads during operation and experiences corresponding tensile and bending stresses along its length.
Blade life is a paramount design objective, and with the introduction of an internal damper, the life of the damper itself affects the life of the blade. The damper should therefore be formed of a material having sufficiently high strength for effecting long low cycle fatigue (LCF) life, long high cycle fatigue (HCF) life, and long rupture life for the damper. These life factors are typically controlled by the highest steady state stress portions of the damper, which is typically in its supporting portion.
In contrast, the outer portion of the damper subject to frictional vibration with the airfoil experiences substantially lower stresses during operation, yet is subject to friction wear.
Standard design practice for introducing blade vibration dampers typically requires a compromise between the wear and strength performance of the damper.
Accordingly, it is desired to provide a turbine blade damper having improved wear resistance in combination with high strength.